Bladed disk assemblies, i.e., discrete blades having dovetails mounted in complementary shaped slots in a rotor disk, are well known in the art. Disk assemblies having integral blades and disks i.e., bl(ade)+integral (d)isk="blisk", are also well known in the art; see, for example, U.S. Pat. No. 4,363,602 to J. R. Martin, entitled "Composite Air Foil and Disk Assembly," and U.S. Pat. No. 4,595,340 to D. D. Klassen et al. entitled "Gas Turbine Bladed Disk Assembly".
The use of a blisk assembly over a bladed disk assembly provides many benefits including increased structural strength and improved aerodynamic performance. In particular, a blisk can be designed for obtaining relatively low radius ratio defined as the inlet root radius divided by the blade tip radius, having values less than about 0.5, and relatively high blade root solidity, defined as the root chord length divided by the distance between adjacent blades, having values greater than about 2.3 for obtaining significant improvements in aerodynamic performance. Blisks also typically include relatively high root slope angles of greater than about 10.degree. since the blisk stage is effective for efficiently compressing airflow in a relatively short axial distance.
Although blisks provide substantial aerodynamic performance benefits, it is deemed desirable to have replaceable blades for more easily repairing any foreign object damage thereto. However, experience has shown that conventional bladed disk assemblies are limited to radius ratios greater than about 0.35-0.5 and solidity less than about 2.2 due to life and strength considerations including low cycle fatigue (LCF) and high cycle fatigue (HCF). It should be appreciated that for any given compressor stage, the number and size of the blades needed for performing the required amount of work is generally a fixed requirement. With this given number of blades, it will be appreciated that for obtaining reduced radius ratios to improve aerodynamic performance, the outer perimeter of the disk must be correspondingly reduced, thusly providing less circumferential space for mounting the blades thereto and thereby increasing solidity.
Accordingly, smaller shank and dovetail portions of the blade are required due to the physical limitations of the decreased circumference for low radius ratio applications. However, inasmuch as the size of the airfoil portion of the blade does not basically change, the required smaller conventional dovetail and shank are structurally inadequate for suitably mounting the blade to the disk. For example, such a conventional shank and dovetail would be relatively more flexible and have less load transfer surface area thus leading to undesirable LCF and HCF life in the dovetail and disk assemblies. In particular, the increased flexibility of a conventional low radius ratio blade would decrease the 2/REV margin in a gas turbine engine. The 2/REV excitation frequency is typical and in order to have acceptable HCF life of the blade, a relatively stiff bladed disk assembly having adequate 2/REV margin is desirable.
Inasmuch as a gas turbine rotor typically operates at substantial rotational speeds, centrifugal force generated by the mass of the rotating blades is substantial. The means for securing the blades to the rotor disk therefore must be able to accommodate the substantial centrifugal forces while obtaining acceptable LCF life and acceptably low axial components of such centrifugal force which would tend to slide the blade axially outward from the disk.